1. Field of the Invention
This invention relates to the field of aircraft flight control systems. More particularly, it pertains to multi-mode aircraft command and response switching between control systems best suited to differing requirements.
2. Prior Art
Modern helicopters are required to perform a wide variety of missions ranging from air-to-air combat at several thousand feet altitude to precision hovering within a few feet of ground in gusty air. The wide range of tasks the helicopter performs requires a control system that can be reconfigured rapidly and smoothly to provide aircraft responses appropriate for a given task. These multi-mode control systems are designed to accommodate requirements for varying command/response characteristics in accordance with flight conditions, speed, altitude, task, and the quality and nature of the visual scene available to the aircraft crew.
The benefit of adjusting the aircraft response characteristics to the mission being performed in the available visual cue environment is well known. Generally, the more degraded the available visual cues, the higher the level of stability required to achieve acceptable handling qualities. Aeronautical design standards require various levels of command/stabilization for varying levels of visual cue environments.
In some cases, changes in aircraft response to account for poor visual cues or to provide improved precision maneuvering capability are made automatically, with reference to groundspeed, airspeed, or some other measurable variable. Usually the pilot is required to initiate these mode changes by engaging a selectable control or stabilization mode such as hover hold or precision flight mode. These control modes are selected for specific purposes including precision flight mode for aerial refueling or shipboard landing, hover hold for external cargo hook-up, and velocity hold for constant airspeed flight.
In current aircraft, discrete paths within the flight control system are included to vary the command/response shaping and provide the pilot with the desired level of command or stability augmentation. An example of this is control law mode switching from a baseline automatic flight control system (AFCS) to a precision flight mode (PFM). Switching between these modes is controlled by a transient free switch, which slowly closes an outer-loop around a lateral axis command model. The outer loop closure causes the command model to produce an attitude-type response as opposed to the rate-type response produced when the outer loop is open. Although this manner of switching provides the required function at the end points before and after the switching dynamics have settled, the command/response characteristics during switching are poorly conditioned and very unpredictable. While the transient free switch is changing state, i.e., neither fully opened nor fully closed, response of the command model to a fixed input is dynamic and damping varies while the transition occurs.
In demonstrator aircraft, lateral axis control laws switch from an attitude command system in low-speed flight to a rate command system in higher speed flight using several transient free switches tuned to provide acceptable intermediate states during the mode transition. However, when command model gains are altered, these transient free switch rates require additional tuning to provide acceptable transient response during mode changes.
In these demonstrator aircraft, the primary flight control system (PFCS) feed-forward shaping, as applied to the longitudinal axis, is switched to an alternate shaping path for AFCS operation. Command through one of these parallel paths is controlled by a transient free switch. While switching between these control systems, the slow transient free switch rate reduces sudden commands caused by the dissimilar AFCS and PFCS feed-forward shaping filters. This implementation may cause liftoff transients, i.e., noncommanded inputs when switching between PFCS and AFCS, if controller input is present.
The main rotor of a helicopter rotatably supports blades having an airfoil shape, which produce aerodynamic lift or thrust as the blades pass through the air. A pitch link attached to each blade changes the angle of attack by applying control force to the blade and rotating it about its pitch axis, thereby affecting the magnitude of lift produced by the rotor. The opposite end of each pitch link is connected to a rotating swashplate, which is connected to a stationary, nonrotating ring located below the rotating ring by bearings, which allow relative rotation of the rings and hold them at the same angle and relative axial position along the rotor shaft. The stationary ring can be raised and lowered along the axis of the rotor shaft, or tilted with respect to that axis by action of control servos or actuators, a longitudinal servo and multiple lateral servos.
To change the angular position of rotor lift, the pitch of each blade is changed individually, i.e., cyclic pitch is applied by causing the longitudinal servo to tilt the rings and main rotor about the rotor shaft. To change the magnitude of rotor lift, the pitch of all blades is change concurrently by raising the rings along the rotor shaft by the same amount, i.e., collective pitch is applied by causing the lateral servos to raise the rings relative to the main rotor.
To prevent a single rotor helicopter from rotating continually about its rotor axis, a tail rotor is used to produce a thrust force directed laterally that compensates for main rotor torque. This stabilizes the yaw heading and attitude of the aircraft against wind gusts and changes in main rotor torque. By overcompensating and undercompensating for these transients, the pilot changes the angular position of the aircraft about the yaw axis.
The magnitude of the tail rotor thrust varies with changes in pitch or angle of attack of the tail rotor blades resulting from raising and lowering a rotating swashplate connected by pitch links to the blades. The position of the swashplate is changed while maintaining its angular position constant so that tail rotor blade pitch changes collectively. Conventionally, the tail rotor thrust is controlled by pilot manipulation of control pedals connected by cables, bellcranks and push-pull rods to the tail rotor controls.